Composite materials are used in a wide variety of structures. In aircraft construction, composite materials may be used to form the fuselage, wings, tail section, and other components. For example, an aircraft fuselage may be constructed of composite skin panels to which composite structural members such as hat stringers may be attached. Hat stringers may increase the strength and stiffness of the skin panels.
During fabrication of a composite structure, layers of composite plies may be laid up over a tool or a mold. The tool or mold may be provided in the desired shape of the final composite structure. The composite plies may comprise a plurality of high-modulus or high-strength fibers such as carbon, glass, or other fibers. The fibers may be pre-impregnated with a polymeric matrix material such as epoxy or thermoplastic resin to form pre-preg composite plies. The fibers in a composite ply may be commonly aligned or oriented in a single direction (e.g., unidirectional) or the fibers in a composite ply may be woven together in two or more directions in a fabric arrangement. Composite structures may be designed to transmit primary loads along the length of the fibers. In this regard, composite structure formed of unidirectional fibers may have a relatively high tensile strength along a lengthwise direction of the fibers.
After the pre-preg composite plies are laid up on the tool or mold, a cure cycle may be performed on the layup. The cure cycle may comprise the application of heat and compaction pressure to the layup. The application of heat may reduce the viscosity of the resin allowing the resin to flow and intermingle with the resin in adjacent composite plies. The application of compaction pressure may include installing a vacuum bag over the layup and/or positioning the layup within an autoclave. The compaction pressure may compact the composite plies against the tool or mold to minimize or reduce porosity and voids in the final composite structure. In addition, the compaction pressure may force the layup against the tool or mold to establish the final shape and surface finish of the composite structure.
Although the vacuum bag may apply substantially uniform pressure to a majority of the layup of pre-preg composite plies, the reduction in resin viscosity during the application of compaction pressure may result in the resin flowing toward regions of low compaction pressure underneath the vacuum bag. The regions of low compaction pressure may occur at locations where there is a geometric discontinuity associated with the layup. The geometric discontinuity may result in out-of-plane fiber movement during curing. For example, a geometric discontinuity may occur at an edge of a structural member (e.g., a stringer, a stiffener, etc.) that may be mounted or joined (co-cured, co-bonded, co-consolidated) to a skin panel formed as a laminate of uncured pre-preg composite plies. The geometric discontinuity at the edge of the stiffener may result in bridging of the vacuum bag from the stiffener edge to the surface of the layup.
The area underneath the bridging may comprise a region of low compaction pressure. Resin may flow toward the region of low compaction pressure and may cause the fibers in the composite plies to also shift toward the region of low compaction pressure. The movement of the fibers may cause the fibers to bunch up resulting in out-of-plane fiber distortion. Upon curing and solidification of the resin, the out-of-plane fiber distortion may become permanently set in the composite structure. The out-of-plane fiber distortion may affect the load-carrying capability of the fibers which are typically designed to provide maximum strength when the fibers are oriented in a common direction within a layer or ply. In this regard, the out-of-plane fiber distortion may have a less than desired effect on characteristics of the final composite structure.
As can be seen, there exists a need in the art for a system and method for minimizing out-of-plane fiber distortion in composite structures.